The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in corresponding turbines which power the compressor, and provide useful work by powering an upstream fan in an exemplary turbofan aircraft engine application.
A high pressure turbine (HPT) directly follows the combustor and receives the hottest combustion gases therefrom. The HPT may have one or more stages therein joined by a shaft to power the compressor.
The LPT typically has several stages following the HPT and is joined by another shaft to the upstream fan in the turbofan application, or instead the shaft may extend externally of the engine for providing output power to drive an electrical generator or transmission in various marine and industrial applications.
Each turbine stage includes a stationary turbine nozzle having a row of stator vanes which direct combustion gases in the downstream direction. A corresponding row of turbine rotor blades follows the nozzle vanes and extracts energy from the combustion gases for in turn rotating the blades on a supporting rotor disk joined to the corresponding drive shaft.
Each nozzle vane has a corresponding crescent or airfoil configuration specifically configured for directing the combustion gases into the downstream row of rotor blades for maximizing energy extraction from the combustion gases. Each vane includes a generally concave pressure side and a circumferentially opposite, generally convex suction side extending axially between corresponding leading and trailing edges, and radially in span between outer and inner supporting bands.
In view of the hostile environment of the combustion gases, the nozzle vanes themselves are typically hollow and joined in flow communication with the compressor for receiving air bled therefrom for use as a coolant in cooling the nozzle vanes during operation against the external thermal loads applied by the hot combustion gases. Any air diverted from the combustion process for cooling the nozzle vanes correspondingly decreases the overall efficiency of the engine, and therefore should be minimized.
The prior art is replete with various configurations for cooling turbine nozzles, which vary in complexity, effectiveness, and in cost of manufacture.
Nozzle materials are typically formed of state-of-the-art nickel-based superalloys which retain strength in the high temperature environment of gas turbines. The superalloy materials nevertheless require suitable cooling during operation for enhancing the useful life and durability of the nozzle during operation.
Further enhancement and nozzle protection in the hostile environment of gas turbines may be achieved by using conventional thermal barrier coatings (TBCs). The TBC is typically a ceramic material which covers the external surfaces of the nozzle-vanes and corresponding flow bounding surfaces of the bands for providing a thermal insulation barrier against the hot combustion gases. The TBC protects the external surfaces of the nozzle vanes, and the internal surfaces thereof may be suitably cooled by the air coolant channeled therethrough during operation.
For example, the hollow nozzle vanes may include impingement inserts or baffles which have suitable patterns of small impingement cooling holes extending therethrough. The baffles are formed of thin superalloy metal, and are spaced from the internal surfaces of the vane sidewalls for permitting the coolant to firstly impinge against those internal surfaces for extracting heat therefrom, with the spent impingement air then being discharged through suitable outlets in the vanes.
Such outlets may include rows of film cooling holes extending through the vane sidewalls, which are typically inclined aft for discharging cooling air in a film that provides an additional thermal barrier or insulation layer between the vane and hot combustion gases. Each vane may also include a row of trailing edge outlet holes which discharge another portion of the spent impingement air through the thin trailing edge for enhanced cooling thereof. And, additional outlet or dump holes may be provided in the supporting bands for discharging additional air therethrough.
The exemplary features described above, among others, increase the sophistication and complexity of manufacturing turbine nozzles, and are necessarily tailored to match the cooling requirements of the different portions of the nozzle against the different thermal loads applied by the combustion gases as they flow with different velocity distributions over the pressure and suction sides of the vanes.
The manufacturing process also affects the design of the nozzle. For example, a typical turbine nozzle is divided into a number of nozzle segments around the perimeter thereof to eliminate the hoop constraint of a unitary ring, and thereby reduce the magnitude of thermal stresses generated during operation. A typical nozzle segment includes a pair of nozzle vanes integrally joined to corresponding arcuate outer and inner bands, with adjoining nozzle segments being sealed together at corresponding axial splitlines by straight spline seals therein. The nozzle segment doublet may be manufactured from constituent parts and then assembled or brazed together, but is typically manufactured in a common casting including the outer and inner band segments and the pair of hollow nozzle vanes.
The impingement baffles are separately manufactured and later installed into corresponding cavities or plenums in the vanes during the assembly process.
The TBC is typically applied using a suitable vapor deposition process to coat the nozzle vanes with a sufficient amount of the TBC material. The film cooling holes may be formed through the nozzle vanes prior to applying the TBC using a suitable drilling process such as electrical discharge machining (EDM). Since the nozzle trailing edge holes are typically formed in the casting process to provide flow communication with the plenums inside the vanes, the trailing edge region of the vanes is preferably masked during the TBC deposition process to prevent clogging of those apertures.
Since the typical nozzle is an annular or axisymmetric assembly, the nozzle segments and vanes are typically identical around the perimeter of the nozzle. Furthermore, the impingement baffles with various patterns of cooling holes in the nozzle vanes are also identical from vane to vane. This therefore limits the number of different parts and drawings required in making the turbine nozzle.
The identical nozzle vanes and their identical cooling configurations therefore ensure substantially identical performance of the turbine nozzle vanes during operation in the engine, with the life or durability of the nozzle being affected by random differences within the manufacturing tolerances of the nozzle parts, and random differences in the distribution of the combustion gases.
However, since the typical TBC vapor deposition process is directional, it is not possible to evenly deposit the TBC over the full external surfaces of the nozzle sidewalls in the doublet configuration. Since the TBC is applied to each nozzle doublet individually, the exposed or outboard surfaces thereof may be readily coated with the TBC to the desired nominal or full thickness thereof, whereas the hidden or inboard surfaces of the nozzle doublet may only be partially coated with a thinner thickness of the TBC.
More specifically, the doublet pair includes a first or leading vane whose convex suction side faces circumferentially outwardly at the corresponding splitline. The second or trailing vane of the doublet has its concave pressure side facing outwardly towards the opposite splitline. The concave pressure side of the leading vane therefore faces circumferentially inwardly toward the opposing convex suction side of the trailing vane, and therefore both of these inboard sidewalls are hidden from the outside of the nozzle by the shadowing effect of their opposite sidewalls in the vanes.
Accordingly, during the TBC vapor deposition process, the trailing vane casts a shadow in the vapor deposition over the inboard pressure side of the leading vane and results in thinner application of the TBC thereon. Correspondingly, the leading vane casts a shadow over the inboard convex suction side of the trailing vane during the TBC vapor deposition process resulting in a correspondingly thin deposition of the TBC thereon.
In contrast, the entire convex suction side of the leading vane faces outboard and may be fully coated with the TBC. And, the entire concave pressure side of the trailing vane faces outboard and may also be fully coated with the TBC. And, the opposite leading and trailing edges also face outboard and may be suitably coated to the desired full thickness.
Since the resulting nozzle doublet coated with TBC in this process would have partial thickness TBC along the pressure side of the leading vane and along the suction side of the trail vane the uniformity or identicality between the two nozzle vanes would be prevented. Correspondingly, cooling performance of the two nozzle vanes would no longer be identical.
Accordingly, conventional practice used in the US for many years introduces suitable masks during the TBC vapor deposition process to effectively create dummy nozzle vanes aligned with the outboard sidewalls of the doublet vanes, typically in the positions of the next adjacent vanes in the fully assembled nozzle ring. In this way, the dummy masks may be used to ensure that the outboard suction side of the lead vane receives partial thickness TBC in the same manner as the inboard suction sidewall of the trail vane.
Correspondingly, the opposite mask ensures that the outboard pressure sidewall of the trail vane receives partial thickness TBC in the same manner as the partial thickness of the TBC on the inboard pressure side of the lead vane.
In this way, the two nozzle vanes in the nozzle doublet segment have substantially identical configurations, and may be similarly cooled during operation using the identical configurations of the impingement baffles and various outlet apertures through the nozzle vanes.
Although the typical nozzle vanes manufactured in accordance with this conventional process therefore have substantially identical cooling system design, the nozzle segments are in fact not subject to identical loading during operation. For example, although the nozzle flow passages between adjacent vanes are substantially identical for channeling the combustion gases therethrough, the circumferential continuity of the nozzle is interrupted by the segment configuration, which in turn affects distribution of the loads in each nozzle segment.
The gas pressure loads are reacted by the nozzle vanes during operation and are carried through the nozzle bands to the corresponding nozzle support. And, the nozzle vanes and their bands are subject to different temperatures during operation which differently expand and contract these components, which in turn leads to differences in thermal loading thereof.
For example, the arcuate outer and inner bands of the nozzle segments are initially aligned in corresponding hoops prior to being heated by the combustion gases. As the gases heat the nozzle segments, the outer band in particular tends to straighten along its chord between the opposite splitlines, which distortion is restrained by the two nozzle vanes attached thereto.
This chording effect introduces additional thermal stress in the inboard sidewalls of the two pressure and suction sides which face each other in the nozzle doublet. And, the outboard sidewalls of the two nozzle vanes defined by the pressure and suction sides exposed at the splitlines experience different thermal loading. The corresponding thermal distortion of the nozzle doublets and the thermal stress introduced thereby adversely affects the durability or useful life of the nozzle segment.
Accordingly, it is desired to provide a turbine nozzle having custom cooling for reducing the adverse effects of the different thermal loading therein.